Field of the Invention
The field of the present invention is that of thermodynamics and more specifically that of the blades for the compressors of turbomachines.
Description of the Related Art
Aeronautical turbomachines are conventionally made up, from upstream to downstream in the direction in which the gases flow, of a fan, of one or more compressor stages, for example a low-pressure compressor and a high-pressure compressor, a combustion chamber, one or more turbine stages, for example a high-pressure turbine and a low-pressure turbine, and a gas exhaust nozzle. The compressor or compressors are produced in the form of a plurality of sets of rotor blading rotating past a plurality of sets of stator blading known as guide vanes. The rotor blading is arranged evenly at the periphery of a disk driven by the rotor of the turbomachine, and their airfoils extend radially between the rotor disk and a casing enclosing the airflow path.
Each rotor blade comprises a pressure face over which the air of the flow path is at a raised pressure with respect to the mean pressure prevailing in the vicinity of the blade airfoil, and a suction face over which the air is at a reduced pressure in relation to this mean pressure. This then causes an air circuit to become established at the outer tip of the blade, causing air to pass from the pressure face to the suction face through the clearance there is between the blade and the casing. In the known way, this circulation of air develops along the entire length of the chord of the blade and takes the form of a vortex, referred to as the blade tip clearance vortex, which spreads downstream of the trailing edge of the blade.
The presence of this vortex disturbs the flow in the stages further downstream of the compressor and creates losses which are detrimental to the efficiency of the compressor. It would therefore be desirable to eliminate this vortex or at the very least, to reduce the flow rate of air it carries.
Attempts have been made to try to control this vortex, these for example including treatments applied to the casing surrounding the compressor or the creation of “trenches”, namely cavities hollowed into the casing. One example of such treatments is described in the applicant's patent application published under the number FR 2940374. All of these have the disadvantage of generating additional cost in producing the turbomachine and of potentially impairing the performance of the compressor in terms of efficiency at certain operating points.
Patent applications have also been filed in an attempt to reduce the impact that this vortex has on the efficiency of a compressor or turbine stage, these including for example applications US 2010/0054946 or EP 1953341. These applications plan to modify the shape of the blades by altering the shape given to the leading edge, i.e. by altering its sweep angle between the root and the tip of the blade along this leading edge. They do not, with the exception of FIG. 12 of the American publication, provide any indication regarding changes to the line of stacking of the elemental profiles along the height of the blade.
Moreover, U.S. Pat. No. 6,341,942 describes undulations along the height of a compressor blade for the purpose of increasing the flexural rigidity thereof, without an increase in its mass. Although it indicates that one undulation may be situated in a position high up on the blade, it does not specify the position of the point of inversion of curvature associated therewith, nor a fortiori does it indicate the position of the lower point of inflection in the case of a double inflection. Moreover, by highlighting the problem of the vibrational behavior of the blade, it is not, a priori, of any benefit to a person skilled in the art wondering how to improve the efficiency of a stage by controlling the blade tip clearance vortex.